U.S. patent number 5,462,405 [Application Number 08/262,396] was granted by the patent office on 1995-10-31 for coolable airfoil structure.
This patent grant is currently assigned to United Technologies Corporation. Invention is credited to Richard W. Hoff, David R. Martin.
United States Patent |
5,462,405 |
Hoff , et al. |
October 31, 1995 |
Coolable airfoil structure
Abstract
A coolable airfoil having internal cooling passages 82, 84, 74
is disclosed. Various construction details are developed to
increase cooling in critical locations of the airfoil. In one
embodiment, a pair of side-by-side spanwisely extending passages
82, 84 flow cooling air to the tip passage 74 which extends
rearwardly in a flag-like manner.
Inventors: |
Hoff; Richard W. (Glastonbury,
CT), Martin; David R. (East Hartford, CT) |
Assignee: |
United Technologies Corporation
(Hartford, CT)
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Family
ID: |
25528753 |
Appl.
No.: |
08/262,396 |
Filed: |
June 20, 1994 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
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981952 |
Nov 24, 1992 |
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Current U.S.
Class: |
416/97R;
415/115 |
Current CPC
Class: |
F01D
5/187 (20130101); F05D 2260/202 (20130101); Y02T
50/60 (20130101); F05D 2260/2212 (20130101); Y02T
50/673 (20130101); Y02T 50/676 (20130101) |
Current International
Class: |
F01D
5/18 (20060101); F01D 005/18 () |
Field of
Search: |
;416/97R,96R,96A
;415/115,116 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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135606 |
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Jul 1985 |
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JP |
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66401 |
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Mar 1989 |
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JP |
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845227 |
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Aug 1960 |
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GB |
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2005775 |
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Apr 1979 |
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GB |
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Other References
Drawing of the first stage turbine blade for the V-2500-A1 engine,
International Aero Engines, 1991..
|
Primary Examiner: Look; Edward K.
Assistant Examiner: Larson; James A.
Attorney, Agent or Firm: Fleischhauer; Gene D.
Parent Case Text
This application is a continuation of application Ser. No.
07/981,952, filed Nov. 28, 1992, abandoned.
Claims
We claim:
1. For a coolable airfoil structure of the type having a leading
edge, a trailing edge, a root section in communication with a
source of cooling air, a first end adjacent the root section, an
airfoil tip region, and a second end at the tip region, the tip
region including a chordwisely extending tip passage extending
rearwardly through the tip region to the trailing edge, and a tip
wall which bounds the tip passage, the improvement which
comprises:
a first spanwisely extending supply passage and a second spanwisely
extending supply passage, a rib which extends spanwisely to bound
the first and second passages in the spanwise direction, each
passage extending from the first end adjacent the root section
toward the second end to the tip region and each independent of the
other being in flow communication with the chordwisely extending
tip passage, the first passage being disposed between the leading
edge and the second passage and being chordwisely forward of the
second passage and separated from the second passage only by the
rib the first passage and second passage extending outwardly to
supply cooling air to the tip region and the rib being spaced from
the tip wall to permit mixing of the cooling air discharged from
the first passage and the second passage;
a plurality of cooling air holes extending from the first
forwardmost passage to the exterior of the airfoil to cool the
surface of the airfoil adjacent the second rearmost passage to
block heat transfer to the second rearmost passage from the
exterior and to discharge a portion of the heated cooling air from
the first forwardmost passage prior to the air mixing with air from
the second passage in the tip passage.
2. A coolable rotor blade for an axial flow rotary machine, the
coolable rotor blade having an exterior, which comprises:
a root section which adapts the rotor blade to engage a rotor
assembly, the root section having a chordwisely extending root
wall, a first duct adapted to be in fluid communication through the
root wall with a source of cooling air, and a second duct adapted
to be in fluid communication through the root wall with a source of
cooling air and, a third duct adapted to be in fluid communication
through the root wall with a source of cooling air;
an airfoil section having,
a leading edge,
a trailing edge,
a suction sidewall,
a pressure sidewall joined to the suction sidewall at the leading
edge and the trailing edge and spaced from the suction sidewall to
form a cavity therebetween,
a tip region having a tip wall extending in a chordwise direction
between the suction sidewall and the pressure sidewall,
a first rib which extends in the spanwise direction to the tip wall
and is spaced from the leading edge to divide the cavity into a
rear portion and a front portion having a first passage,
a second rib which extends in the spanwise direction and is spaced
chordwisely from the first rib to divide the rear portion of the
cavity into a trailing edge region and a midchord region and which
extends in the chordwise direction and is spaced spanwisely from
the tip wall leaving a tip passage in flow communication with the
midchord region,
a third rib which is spaced from the tip wall and extends in the
spanwise direction to divide the midchord region of the blade into
a second passage and a third passage which each extend spanwisely
outwardly away from the root section toward the tip wall, the
second and third passages each being bounded by the third rib and
each being in flow communication with the second duct such that the
cooling air for each mid-chord region passage does not pass through
the other mid-chord region passage and each being in flow
communications with the tip passage, wherein a plurality of cooling
air holes extend through the pressure sidewall and the suction
sidewall to place the second passage in flow communication with the
exterior of the blade and to duct cooling air over the exterior of
the blade and in the direction of the trailing edge under operative
conditions to block the transfer of heat from the exterior to
cooling air flowing in the third passage to increase the cooling
effectiveness of the air in the third passage in the tip region of
the airfoil and wherein the spacing of the third rib from the tip
wall permits the mixing of the cooling air discharged from the
second passage and the third passage.
3. The coolable rotor blade of claim 2, which further has a
trailing edge circle at the trailing edge which is tangent to the
suction sidewall and pressure sidewall and wherein the pressure
sidewall surface in the trailing edge region is substantially
planar such that the pressure sidewall surface is workable with a
fiat grinding surface and wherein the suction sidewall diverges in
the spanwisely outward direction from the pressure sidewall to
thicken the tip region without affecting the planar surface of the
pressure sidewall but is of a thickness such that the tip passage
is bounded by the pressure sidewall and the suction sidewall at
least at a location which is spaced forwardly from the trailing
edge by distance which is equal to approximately one-half the
diameter of a the trailing edge circle which is tangent at the
rearmost portion of the trailing edge to the suction sidewall and
the pressure sidewall.
4. The coolable rotor blade of claim 2 wherein the cooling air
holes extend through the suction sidewall over at least the
outermost fifty (50%) percent of the span of the suction sidewall
of the airfoil and over at least the outermost eighty (80%) percent
of the span of the pressure sidewall of the airfoil.
5. The coolable rotor blade of claim 4 wherein the cooling air
holes extend through the suction sidewall over at least the
outermost sixty (60%) percent span of the suction sidewall of the
airfoil and over the outermost ninety percent span of the pressure
sidewall of the airfoil.
6. The coolable rotor blade for an axial flow rotary machine of
claim 2 wherein the third passage is in the middle third portion of
the air foil as measured in the chordwise direction.
Description
TECHNICAL FIELD
This invention relates to coolable airfoil structures of the type
used in high temperature rotary machines, and more specifically, to
structure for providing cooling fluid to a critical location of the
airfoil. The concepts disclosed have application to both turbine
vanes and turbine blades.
BACKGROUND ART
An axial flow rotary machine, such as a gas turbine engine for an
aircraft, includes a compression section, a combustion section and
a turbine section. A flow path for hot working medium gases extends
axially through the engine. The flow path for hot gases is
generally annular in shape.
As working medium gases are flowed along the flow path, the gases
are compressed in the compression section causing the temperature
and pressure of the gases to rise. The hot, pressurized gases are
burned with fuel in the combustion section to add energy to the
gases. These gases are expanded through the turbine section to
produce useful work and thrust.
The engine has a rotor assembly in the turbine section which is
adapted by a rotor disk and blades extending outwardly therefrom to
receive energy from the hot working medium gases. The rotor
assembly extends to the compression section. The rotor assembly has
compressor blades extending outwardly across the working medium
flow path. The high-energy working medium gases in the turbine
section are expanded through the turbine blades to drive the rotor
assembly about its axis of rotation. The compressor blades rotate
with the rotor assembly and drive the incoming working medium gases
rearwardly, compressing the gases and imparting a swirl velocity to
the gases.
Each rotor blade has an airfoil to direct the hot working medium
gases through the stage of rotor blades and to receive work from
the gases. As a result, the airfoils are bathed in hot working
medium gases during operation causing thermal stresses in the
airfoils. These thermal stresses affect the structural integrity
and fatigue life of the airfoil. In addition, rotational forces
acting on the rotor blade as the rotor blade is driven about the
axis of rotation further increase the stresses to which the blade
is subjected.
Rotor blades are typically cooled to reduce thermal stresses and
thereby provide the rotor blade with a satisfactory structural
integrity and fatigue life.
An example of such a rotor blade is shown in U.S. Pat. No.
4,474,532 entitled "Coolable Airfoil For a Rotary Machine", issued
to Pazder and assigned to the assignee of this application. Another
example of a coolable rotor blade is shown in U.S. Pat. No.
4,278,400 issued to Yamarik and Levengood entitled "Coolable Rotor
Blade" and assigned to the assignee of this application. Each of
these rotor blades is provided with a plurality of cooling air
passages on the interior of the blade. Cooling air is flowed
through the passages to the rearmost portion of the rotor blade,
commonly referred to as the trailing edge, from whence the cooling
air is exhausted into the working medium flow path.
The above art notwithstanding, scientists and engineers working
under the direction of applicant's assignee are seeking to develop
coolable airfoils for use in high temperature rotary machines which
have acceptable level of stresses in critical regions of the
airfoil.
DISCLOSURE OF INVENTION
According to the present invention, a coolable airfoil having a
passage for cooling one end of the airfoil is supplied with cooling
air from the other end of the airfoil via two spanwisely extending
passages with the forward passage supplying film cooling air to the
exterior of the airfoil to shelter the rear passage from the hot
working medium gases on the interior of the engine.
In accordance with one detailed embodiment of the invention, the
tip passage extends rearwardly to the trailing edge of the airfoil
and is bounded by both the pressure sidewall and the suction
sidewall in the trailing edge region of the blade.
In accordance with one detailed embodiment of the invention, the
pressure sidewall is substantially planar over the entire length of
the airfoil and the suction sidewall diverges from the pressure
sidewall in the tip region of the airfoil to provide a thickened
tip to the airfoil.
A primary feature of the present invention is a coolable airfoil
having a pair of passages extending spanwisely from one end of the
airfoil to the other. The passages are in flow communication with a
source of cooling air. Another feature is that the forward most
passage is in flow communication with the exterior of the airfoil
via film cooling holes. The holes extend over at least 50% of the
span of the airfoil on at least one of the sidewalls. In one
detailed embodiment, the passages are in flow communication with a
tip passage. The tip passage extends to the trailing edge region.
Another feature is a suction sidewall and a pressure sidewall which
bound the tip passage to the trailing edge region of the airfoil.
In one detailed embodiment, the tip of the airfoil is thicker in
the circumferential direction than in the remainder of the
airfoil.
A primary advantage of the present invention is the level of
fatigue life which results from cooling the tip region with cool
air provided from a sheltered cooling air supply passage. Another
advantage is the size of the tip which permits the deposition of
abrasive particles of varying sizes and mounts as compared with
airfoils having less thick tip regions. Another advantage is the
cooling effectiveness which results from having the sidewalls
extend to the trailing edge to shelter the interior of the airfoil
from the hot working medium gases.
Other features and advantages will be apparent from the
specification in claims and from the accompanying drawings which
illustrate an embodiment of the invention.
BRIEF DESCRIPTION OF DRAWINGS
FIG. 1 as a side elevation view of an airfoil, such as a rotor
blade, partly in section and partly broken away to show the suction
sidewall of the interior of the airfoil;
FIG. 2 is a cross-sectional view taken along the line 2--2 of FIG.
1;
FIG. 2A is an enlarged view of a portion of FIG. 2;
FIG. 3 is a diagrammatic view taken along the lines 3--3 of FIG. 1
showing the planar pressure surface and the diverging suction
sidewall of the airfoil;
FIG. 4 is a cross-sectional view of tip of the airfoil showing the
relationship of the pressure sidewall to the suction sidewall. And,
in particular, illustrating the nearly parallel nature of the two
sidewalls and the transition region which extends between the
sidewalls;
FIG. 4a is an enlarged view of a portion of FIG. 4.
FIG. 5 is a view taken in the direction 5--5 of FIG. 2 to show the
relationship of the sidewalls and with a portion of the tip broken
away;
FIG. 6 is a view taken along the lines 6--6 FIG. 2 to show the
relationship of the sidewalls and with a portion of the tip and the
pedestals broken away for clarity.
BEST MODE FOR CARRYING OUT THE INVENTION
FIG. 1 shows an airfoil structure such as a rotor blade 10, for a
rotary machine. The airfoil structure includes an airfoil section
12 having a first or innermost end 14 and a second or outermost end
16. The rotor blade has a spanwise reference direction S and
chordwise reference direction C. The airfoil has a length (span)
D.sub.S between its innermost end and its outermost end. As will be
realized, the airfoil structure might be a stator vane.
The rotor blade has a root section 18 and platform section 20
adjacent to the first end. The root section is adapted to engage
the rotor of a rotary machine.
The root section has a chordwisely extending root wall 22. A first
duct 24, a second duct 26, and a third duct 28 are in flow
communication with a source of cooling air, such as a compressor
(not shown). The platform section is adapted to form a portion of
the inner wall of the flow path for working medium gases in a
rotary machine. The airfoil section 12 is adapted to extend
outwardly across the working medium flow path and has a tip region
32 at its most outward (outermost) end.
The airfoil section has a leading edge 34 and a trailing edge 36. A
suction sidewall 38 and a pressure sidewall 42 (partially broken
away in FIG. 1 for clarity and shown in FIG. 2) are joined at the
leading edge and the trailing edge. The pressure sidewall is spaced
from the suction sidewall to form a cavity 44 therebetween. An
internal tip wall 46 extends between the pressure sidewall and the
suction sidewall to bound the cavity in the spanwise direction. An
internal wall, such as a first rib 48, extends in the spanwise
direction to the tip wall and is spaced from the leading edge to
divide the cavity into a rear portion 52 and a front portion 54.
The front portion has at least one passage as represented by the
first passage 56. In the embodiment shown, the leading edge has a
plurality of impingement spaces 58 disposed in the first passage. A
rib 62 is disposed in the first passage to bound the impingement
spaces. The rib has impingement holes 64 extending therethrough
which supply the impingement spaces to provide an impingement
cooling to the leading edge.
The rear portion 52 has a second rib 66. The second rib extends in
the spanwise direction and is spaced chordwisely from the first rib
48. The second rib also divides the rear portion of the cavity into
a trailing edge region 72 and a midchord region 68. The second rib
also extends in the chordwise direction and is spaced spanwisely
from the tip wall leaving a tip passage 74 in flow communication
with the mid-chord region 68. Cooling air holes 76 extend from the
tip region to the exterior.
The rear portion has a third rib 78. The third rib is spaced from
the tip wall. The third rib extends in the spanwise direction to
divide the mid-chord region of the blade into a second passage 82
and a third passage 84 which each extends spanwisely from the mot
region 18. Each passage is in flow communication with the second
duct 26 and in flow communication with the tip passage 74. A
plurality of film cooling holes 86 are in flow communication with
the second passage. The holes extend from the interior of the
airfoil to the exterior of the airfoil. The holes extend over at
least the outermost fifty (50%) percent of the span of at lest one
of the sidewalls. In the embodiment shown the holes extend over 60%
of the span on the suction sidewall surface and over 90 % of the
span on the pressure sidewall adjacent the second passage. The
holes on the pressure side surface are not shown because the
pressure sidewall is broken away for clarity. In another
embodiment, the holes extend through the suction sidewall over at
least the outermost fifth (50%) percent of the span of the airfoil
and over at least the outermost eighty (80%) percent of the span of
the airfoil. As noted above, the holes extend through the pressure
sidewall.
The rear portion 52 has a fourth rib 80. The fourth rib extends
spanwisely from the tip wall 46 to the root region 18 and is spaced
from the root wall 22. The fourth rib bounds the second passage 82
and is spaced from the first rib leaving a fourth radially
extending passage 92 therebetween. A spanwisely extending supply
passage 94 extends from the second duct 26 to the second and fourth
passages 82, 92 to place the second and fourth passages in flow
communication with the second duct. A plurality of radial holes 95
in the tip region place the passages 58, 92,82, 84, and 74 in flow
communication with the exterior of the blade.
The trailing edge region has a plurality of slots 96 at the rear of
the airfoil. The slots are bounded by tear-drop shaped pedestals or
lands 98 which extend from the pressure sidewall 42 to the suction
sidewall. The trailing edge region includes a fifth spanwisely
extending 102 passage. The fifth passage is space chordwisely from
the slots for ducting cooling air into the trailing edge region. A
pair of spanwisely extending ribs 104, 106 are disposed between the
fifth supply passage and the slots. Each has a plurality of
impingement 108 holes extending therethrough to direct cooling air
against adjacent structure. For example, the holes in the fast rib
direct cooling air against the second rib. The holes in the second
rib direct cooling air against the pedestals.
FIG. 2 is an airfoil cross section taken along the lines 2--2 of
FIG. 1 at approximately the 65% span location of the airfoil. The
airfoil section is partially broken away and partly in section for
city. The cooling air holes 86 extending through the second passage
84 direct cooling air outwardly and rearwardly as shown by the
arrows 88. The suction sidewall 38 and the planar pressure sidewall
42 extend rearwardly. The pressure sidewall is cut away to form a
thin trailing edge region 72 of the airfoil.
FIG. 3 is a diagrammatic representation of the airfoil 12 looking
forwardly towards the trailing edge region 72 of the airfoil. The
pressure sidewall is nearly planar along the entire length of the
span. The pressure sidewall is planar to the extent needed to use a
flat grinding device to remove material during manufacturing. The
suction sidewall diverges from the pressure sidewall in the tip
region of the airfoil. This causes the tip region to be much
thicker in the circumferential direction than the spanwisely
innermost portion of the airfoil. The airfoil tapers outwardly to
its thickest portion.
FIG. 4 is a cross-sectional view taken along the line 4--4 of FIG.
1 and parallel to the sectional plane 2--2 shown in FIG. 1. As
shown in FIG. 4, the pressure sidewall 42 and the suction sidewall
38 extend rearwardly side by side and nearly parallel to the nearly
most rearward portion of the airfoil. In this region of the
airfoil, a curved transition surface 110 (normally circular in
cross-section) connects the suction sidewall to the pressure
sidewall. Both the suction sidewall and the pressure sidewall are
tangential to the transition surface circle. As shown in FIG. 4a,
the diameter of the circular surface is approximately ninety (90)
mils and is much smaller as shown in FIG. 2a. In a typical airfoil
the smaller diameter as shown in FIG. 2 is about thirty (30)
mils.
FIG. 5 and FIG. 6 are sectional views taken along the line 5--5 and
line 6--6 of FIG. 2. FIG. 5 is taken at a distance which is less
than the diameter of the transition surface of the rearward most
portion of the airfoil. As can be seen, the suction sidewall 38 and
the pressure sidewall 42 bound the tip passage all the way to the
rear of the pressure sidewall and suction sidewall.
During operation of the rotary machine, hot working medium gases
are flowed over the exterior surface of the airfoil section. Heat
is transferred from the gases to the suction sidewall 38 and the
pressure sidewall 42. Cooling air is flowed from the first duct 24
via the first passage 56 in the radial spanwise direction and
thence through the tip holes 95 through the tip of the airfoil.
Additional amounts of cooling air are flowed via the impingement
holes 64 in the impingement ribs 62 into the impingement spaces 58.
The cooling air is impinged against the leading edge of the airfoil
to cool the leading edge.
Cooling air is flowed from the second duct 26 via the supply
passage 94 and directly from the second duct via the third 84
passage to the tip passage. The cooling air is then flowed
chordwisely to the rear of the airfoil to cool the tip region. The
tip region has a very high heat load because heat transfer is not
only from the sidewalls. Heat is also transferred through the tip
of the airfoil unit is also through the tip into the tip region of
the airfoil. As cooling air is flowed radially outwardly in the
second passage 82, cooling air is flowed to the exterior via the
film cooling holes 86. These holes in the pressure sidewall and the
suction sidewall provide film cooling to the exterior of the
airfoil which cools the airfoil and which also blocks heat transfer
from the hot gases of the working medium flow path to the air in
the third passage. As a result the air in the third passage is
relatively cool compared to the air in the second passage. The
third passage air is much cooler than if it were not film cooled on
its way to the tip of the airfoil.
A particular advantage of this construction is that the hotter air
in the second passage is used for two purposes: 1) to film cool the
airfoil; and 2) to prevent excessive heat transfer to the air in
the third passage as it moves to the tip region. A second advantage
of this construction is that the hotter air in the second passage
is vented overboard prior to mixing with the cooler air from the
third passage in the tip passage, again to provide the tip passage
with lower temperature cooling air. The lower temperature cooling
air is very effective for providing cooling to the tip of the
airfoil.
Another advantage of the location of the chordwisely extending
cooling air passage is that the supply passages 82, 84 are in the
middle third of the air foil. The third passage 84 is located in a
region which has a lower heat load than does the second passage.
The higher heat load of the second passage is absorbed by the
cooling air in the second passage and that heat energy is dumped
overboard to preserve the cooler temperature of the air supplied by
the third passage to the tip passage.
As the air in the tip passage 74 flows chordwisely rearwardly, it
removes heat from the suction sidewall 38 and the pressure sidewall
42 which are shielding the interior of the airfoil. The shielding
is not removed by a cutaway pressure surface by reason of the
thickened tip which permits both a tip passage of adequate cross
section to flow the cooling air out of the airfoil while providing
a pressure sidewall and suction sidewall to shield the interior.
Another benefit is that the radially extending cooling air passages
in the tip may be extended further back in the tip region to
provide cooling to this critical location of the tip.
The planar pressure sidewall 42 in the rearmost trailing edge
region of the airfoil provides for ease of fabrication by grinding,
for example, with a belt sander during finishing operation of the
casting. If a non-linear trailing edge were used, grinding the
trailing edge with a flat belt would tend to remove metal in
critical locations that can result in an unacceptably thin wall in
those areas.
The thicker tip region 32 of the airfoil avoids the difficulty in
thin airfoils of installing a cooling air hole in one surface which
might penetrate the other surface. Secondly, the thicker trailing
edge tip region allows for a larger cross-sectional area at the
exit of the tip passage. This prevents high local cooling air
velocities which can cause a corresponding reduction in the local
static pressure in the cooling air passage. This avoids an internal
static pressure which drops below the static pressure of the gas
path and maintains the necessary margin of safety against back
flow. As will be realized, if the hot working medium gases flow
into the airfoil because of the negative static pressure gradient,
severe heating and possible cracking of the trailing edge region
will result.
The thicker trailing edge region 32 also avoids the complication
that occurs when installing abrasive grits on the blade tip. These
grits are provided so that the turbine blade can rub acceptably
against an abradable mating seal. The grit size is customarily set
by considerations other than turbine aerodynamics or durability.
For example, the grit size is the grit size picked to adequately
rub the outer airseal. For a given grit size, the number of grits
must necessarily decrease as the thickness of the trailing edge
decreases. Thus, with thin trailing edges the designer may need to
pick between the optimum grit count or the optimum grit size. This
is avoided by using the thicker trailing edge.
Although this invention has been shown and described with respect
to a preferred embodiment, it will be understood by those skilled
in this art that various changes in form and detail thereof may be
made without departing from the spirit and scope of the claimed
invention.
* * * * *